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US3173251A - Apparatus for igniting solid propellants - Google Patents

Apparatus for igniting solid propellants Download PDF

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US3173251A
US3173251A US180370A US18037062A US3173251A US 3173251 A US3173251 A US 3173251A US 180370 A US180370 A US 180370A US 18037062 A US18037062 A US 18037062A US 3173251 A US3173251 A US 3173251A
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fluid
tube
solid propellant
bore
hypergolic
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US180370A
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Jr Harrison Allen
Carl C Ciepluch
Fletcher Edward Abraham
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42C—AMMUNITION FUZES; ARMING OR SAFETY MEANS THEREFOR
    • F42C19/00—Details of fuzes
    • F42C19/08—Primers; Detonators
    • F42C19/0819—Primers or igniters for the initiation of rocket motors, i.e. pyrotechnical aspects thereof
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements

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  • This invention pertains to the ignition of solid propellant rocket motors and, more particularly, to the ignition of solid propellant rocket motors by the injection of hypergolic fluids.
  • Solid propellant rockets are customarily ignited by firing a small quantity of a pyrotechnic powder, such as black powder, in the form of a squib.
  • Igniters of this type have many disadvantages, such as great temperature sensitivity which requires certain environmental conditions, unpredictable shelf life which results in deterioration, complicated mechanical parts which may not be reliable, and rigid quality control requirements.
  • an object of the invention to provide an improved apparatus for igniting a mass of solid propellants using a hypergolic igniting fluid.
  • Another object of the invention is to provide an apparatus for igniting a mass of solid propellant utilizing hypergolic fluids wherein the igniting substance is concentrated on predetermined portions of the propellant.
  • a still further object of the invention is to provide an apparatus for studying the ignition of a solid propellant rocket using a hypergolic fluid.
  • FIG. 1 is a partially sectioned, perspective view of a composite propellant motor having improved ignition apparatus for performing the method of the invention
  • FIG. 2 is an enlarged sectional view of the preferred embodiment of a hypergolic fluid injector constructed in accordance with the invention
  • FIG. 3 is an enlarged sectional view of an alternate embodiment of the injector
  • FIG. 4 is an enlarged sectional view of another embodiment of the injector.
  • FIG. 5 is an enlarged sectional view of still another embodiment of the injector.
  • a method of igniting a tubular mass of solid propellant in which a measured quantity of hypergolic fluid is injected into an opening in the tubular mass, and this fluid is directed against predetermined portions of the walls of the opening.
  • apparatus for performing the method of the invention which includes an injector mounted at one end of a propellant charge for supplying the hypergolic fluid into the interior 3,173,251 Patented Mar. 16, 1965 of the tube, and means for concentrating the fluid at the tube end to react with the propellant before escaping from the opposite end of the rocket nozzle.
  • FIG. 1 a testing apparatus for determining the effects of different ignition fluids and various injector configurations, together with other features relating to the ignition of a rocket engine.
  • This testing apparatus comprises a solid propellant motor 11 having a head 12 such as a block of steel or other similar material located at its upstream end.
  • a tap 13 for conveying an ignition fluid and a chamber pressure passage 14 are drilled in the block 12.
  • the testing apparatus utilizes an ignition system which includes a storage bottle 15 connected to a control valve 16 that may be operated remotely by a powered mechanical device (not shown).
  • the bottle 15 is previously filled with a measured quantity of ignition fluid and pressurized with a non-reactive gas such as helium.
  • the ignition system has a check valve 17 between the control valve 16 and the tap 13.
  • a double-Walled cylindrical chamber 18 encloses a hollow charge consisting of a mass of solid propellant in the form of an elongated tube 19.
  • the solid propellant in the tube 19 is preferably a composite consisting primarily of polysulfide polymer fuel and ammonium perchlorate oxidizer.
  • the chamber 18 is rigidly attached to a protuberance in the form of a threaded boss 20 extending from a surface of the head 12.
  • An opening comprising a bore having a plurality of longitudinal grooves 21 formed in the wall thereof extends along the entire axial length of the tube 19 from a nozzle 22 to the boss 20.
  • the nozzle 22 is of the conventional convergentdivergent type and is rigidly afiixed to the downstream end of the cylindrical casing 18 for conducting gases exhausted from the rocket engine 10.
  • hypergolic ignition fluid such as chlorine trifluoride, bromine trifluoride, or bromine pentafluoride stored in the bottle 15 in a measured quantity such as five cubic centimeters, passes through the valves 16 and 17 into the tap 13. This measured quantity of fluid is then introduced into the head end of .the opening in the propellant charge 19 by a suitable injector apparatus 23 associated with the boss 20 where the fluid is directed against preselected portions of the inside walls of the opening.
  • a pressure transducer 24 is positioned adjacent the head 12 by an intermediate block 25 through which the passage 14 in the head 12 also passes.
  • the transducer 24 is connected to the passage 14 for measuring the chamber pressure, and this device can be of the strain gage type.
  • the pressure side of the transducer 24 and vertical pressure line above it are filled with a high temperature silicone oil for protection from combustion gases.
  • An electrical signal from the pressure transducer 24 is recorded on a direct recording oscillograph 26 which also measures ignition delay. This is accomplished by imposing a separate signal on the oscillograph record at the moment when the valve 16 is opened thereby marking time zero. The ignition delay is the elapsed time from the marked time zero until the chamber pressure rises to 50 percent of its normal value.
  • This ignition delay comprises both dead time and reaction time; the dead time being the time required for the fluid to reach the injector after the opening of the valve 16 while the reaction time is determined by subtracting the dead time from the ignition delay.
  • the dead time which is determined by high-speed photography is a constant of about 1.1 seconds for the disclosed testing apparatus.
  • the preferred form of the invention is illustrated by the pocket type injector 23 in FIGS. 1 and 2.
  • This injector comprises an elongated member such as a tube 30 that extends into the bore formed in the mass of solid propellant 19.
  • a longitudinally extending passage in the tube 30 is in communication with the tap 13 for conveying hypergolic fluid into an enlarged pocket 31 formed in the end of the charge of solid propellant 19.
  • a mass of solid propellant such as a disk 32 of plasticized polyvinyl chloride and ammonium perchlorate is positioned adjacent the tube 30 at the bottom of the pocket 31. The fluid is concentrated in the pocket 31 to react with both the solid propellant of the tube 19 and the disk 32 before it escapes through the nozzle 22 after the disk 32 burns.
  • the pocket type injector 23 was mounted in the apparatus shown in FIG. 1, and two tests were performed using cubic centimeters of chlorine trifluoride pressurized with 15 p.s.i.g. helium gas.
  • the tube 30 had an inside diameter of 0.104 inch and extended into a one inch diameter pocket 31 for a distance of /a-inch.
  • the disk 32 was positioned 1% inches from the head end of the charge 19, and the two tests produced reaction times of 0.31 and 0.34 second.
  • the injector means 23a comprises a longitudinally extending tube 35 having its innermost end 36 closed.
  • a plurality of radial apertures 37 extend through the walls of the tube 35 to concentrate the hypergolic fluid. on the walls of the solid propellant charge 19.
  • a radial injector of the type shown in FIG. 3 comprising a tube 35 having an inside diameter of 0.104 inch I was tested in the apparatus shown in FIG. 1.
  • This tube 35 extended into the bore for a distance of A-inch and ⁇ four radially-drilled holes 37, each having a diameter of 0.06 inch were provided for directing the fluid against the solid propellant.
  • Hypergolic fluid in the storage bot- I tle 15 was pressurized to 15 p.s.i.g. and reaction times of 0.55 and 0.59 second were observed.
  • hypergolic fluid is conveyed to the head end of the tube 19 of solid propellant through the tap 13 as apparatus shown in FIG. 1.
  • This injector included a pair 1 of drilled passages each having a diameter of 0.072 inch that were located at an angle of 90 degrees from one another. A spacing of 0.313 inch was provided between the ends of the passages 40 and 41, and a reaction time of 0.71 second was observed when the hypergolic fluid in the bottle 15 was pressurized to 15 p.s.i.g.
  • the axial type injector 230 shown in FIG. 5 consists of an elongated tube 44 coaxial with the bore in the tube 19 of solid propellant.
  • a tube 44 having an inside diameter of 0.104 inch was mounted on the apparatus shown in FIG. 1 to extend into the bore in the solid propellant charge 19, a distance of /2-inch.
  • a reaction time of 0.91 second was observed when the hypergolic fluid was pressurized to 15 p.s.i.g.
  • the invention is not limited to any particular type of solid propellant, and the various types of injectors that have been disclosed as being used with composite solid propellants are equally applicable to double-base propellant motors.
  • the hypergolic fluids are not limited to the three particular aforementioned fluorides.
  • Apparatus for igniting a tube of solid propellant having a longitudinal bore extending therethrough comprising means at one end of said tube for directing a stream of hypergolic fluid along the axis of said bore, and
  • Apparatus for igniting a tube of solid propellant having a longitudinal bore extending therethrough comprising injector means mounted at one end of said tube for introducing hypergolic fluid into said bore, said injector means comprising an elongated member extending into said bore at said one end thereof, said member having a longitudinally extending passage therein for conveying said fluid therethrough into said bore, and
  • a mass of solid propellant positioned in said bore adjacent said elongated member for forming an obstruction which initially blocks the flow of said hypergolic fluid thereby concentrating said fluid at said one end to react with said solid propellant in said tube and said mass before escaping from the opposite end of said bore.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • Testing Of Engines (AREA)

Description

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March 16, 1965 H. ALLEN, JR., ETAL 3,173,251
APPARATUS FOR IGNITING soup PROPELLANTS Filed March 16, 1962 2 Sheets-Sheet 1 INVENTORS HARRISON ALLEN JR. CARL C. CIEPLUGH EDWARD A.FLETCHER jlfl$flg ATTORNEY March 16, 1965 H. ALLEN, JR.. ETAL 3,173,251
APPARATUS FOR IGNITING soun PROPELLANTS Fileg March 16, 1962 2 Sheets-Sheet 2 FIG.'3
FIG. 5
INVENTOR HARRISON ALLEN JR.
CARL 0' CIEPLUCH EDWARD A.FLETCHER BY jfl0%v;
ATTORNEY United States Patent 3,173,251 APPARATUS FOR IGNITING SOLID PROPELLANTS Harrison Allen, Jr., Cleveland, and Carl C. Ciepluch,
Berea, Ohio, and Edward Abraham Fletcher, Minneapolis, Minn., assignors t0 the United States of America as represented by the Administrator of the National Aeronautics and Space Administration Filed Mar. 16, 1962, Ser. No. 180,370 4 Claims. (Cl. 60--35.6) (Granted under Title 35, US. Code (1952), see. 266) The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.
This invention pertains to the ignition of solid propellant rocket motors and, more particularly, to the ignition of solid propellant rocket motors by the injection of hypergolic fluids.
Solid propellant rockets are customarily ignited by firing a small quantity of a pyrotechnic powder, such as black powder, in the form of a squib. Igniters of this type have many disadvantages, such as great temperature sensitivity which requires certain environmental conditions, unpredictable shelf life which results in deterioration, complicated mechanical parts which may not be reliable, and rigid quality control requirements.
In order to solve these problems, it has been proposed to distribute a chemical fluid such as bromine trifluoride or chlorine trifluoride over a large area of a mass of solid propellant. The fluid reacts hypergolically with either the fuel or oxygen component of the propellant grains to cause ignition. However, this ignition is not simultaneous and an objectionable delay results which makes firing at a precise time difiicult to obtain.
It is, therefore, an object of the invention to provide an improved apparatus for igniting a mass of solid propellants using a hypergolic igniting fluid.
Another object of the invention is to provide an apparatus for igniting a mass of solid propellant utilizing hypergolic fluids wherein the igniting substance is concentrated on predetermined portions of the propellant.
A still further object of the invention is to provide an apparatus for studying the ignition of a solid propellant rocket using a hypergolic fluid.
Other objects of the invention will be apparent from the following description when taken together with the accompanying drawings wherein like numerals are used throughout to identify like parts.
In the drawings:
FIG. 1 is a partially sectioned, perspective view of a composite propellant motor having improved ignition apparatus for performing the method of the invention;
FIG. 2 is an enlarged sectional view of the preferred embodiment of a hypergolic fluid injector constructed in accordance with the invention;
FIG. 3 is an enlarged sectional view of an alternate embodiment of the injector;
FIG. 4 is an enlarged sectional view of another embodiment of the injector; and
FIG. 5 is an enlarged sectional view of still another embodiment of the injector.
In order to achieve the above listed objects there is provided a method of igniting a tubular mass of solid propellant in which a measured quantity of hypergolic fluid is injected into an opening in the tubular mass, and this fluid is directed against predetermined portions of the walls of the opening. There is also provided apparatus for performing the method of the invention which includes an injector mounted at one end of a propellant charge for supplying the hypergolic fluid into the interior 3,173,251 Patented Mar. 16, 1965 of the tube, and means for concentrating the fluid at the tube end to react with the propellant before escaping from the opposite end of the rocket nozzle.
Referring to the drawings, there is shown in FIG. 1 a testing apparatus for determining the effects of different ignition fluids and various injector configurations, together with other features relating to the ignition of a rocket engine. This testing apparatus comprises a solid propellant motor 11 having a head 12 such as a block of steel or other similar material located at its upstream end. A tap 13 for conveying an ignition fluid and a chamber pressure passage 14 are drilled in the block 12. The testing apparatus utilizes an ignition system which includes a storage bottle 15 connected to a control valve 16 that may be operated remotely by a powered mechanical device (not shown). The bottle 15 is previously filled with a measured quantity of ignition fluid and pressurized with a non-reactive gas such as helium. Additionally, the ignition system has a check valve 17 between the control valve 16 and the tap 13.
A double-Walled cylindrical chamber 18 encloses a hollow charge consisting of a mass of solid propellant in the form of an elongated tube 19. The solid propellant in the tube 19 is preferably a composite consisting primarily of polysulfide polymer fuel and ammonium perchlorate oxidizer. The chamber 18 is rigidly attached to a protuberance in the form of a threaded boss 20 extending from a surface of the head 12. An opening comprising a bore having a plurality of longitudinal grooves 21 formed in the wall thereof extends along the entire axial length of the tube 19 from a nozzle 22 to the boss 20. The nozzle 22 is of the conventional convergentdivergent type and is rigidly afiixed to the downstream end of the cylindrical casing 18 for conducting gases exhausted from the rocket engine 10.
According to the invention, hypergolic ignition fluid, such as chlorine trifluoride, bromine trifluoride, or bromine pentafluoride stored in the bottle 15 in a measured quantity such as five cubic centimeters, passes through the valves 16 and 17 into the tap 13. This measured quantity of fluid is then introduced into the head end of .the opening in the propellant charge 19 by a suitable injector apparatus 23 associated with the boss 20 where the fluid is directed against preselected portions of the inside walls of the opening.
A pressure transducer 24 is positioned adjacent the head 12 by an intermediate block 25 through which the passage 14 in the head 12 also passes. The transducer 24 is connected to the passage 14 for measuring the chamber pressure, and this device can be of the strain gage type. The pressure side of the transducer 24 and vertical pressure line above it are filled with a high temperature silicone oil for protection from combustion gases. An electrical signal from the pressure transducer 24 is recorded on a direct recording oscillograph 26 which also measures ignition delay. This is accomplished by imposing a separate signal on the oscillograph record at the moment when the valve 16 is opened thereby marking time zero. The ignition delay is the elapsed time from the marked time zero until the chamber pressure rises to 50 percent of its normal value. This ignition delay comprises both dead time and reaction time; the dead time being the time required for the fluid to reach the injector after the opening of the valve 16 while the reaction time is determined by subtracting the dead time from the ignition delay. The dead time which is determined by high-speed photography is a constant of about 1.1 seconds for the disclosed testing apparatus.
The preferred form of the invention is illustrated by the pocket type injector 23 in FIGS. 1 and 2. This injector comprises an elongated member such as a tube 30 that extends into the bore formed in the mass of solid propellant 19. A longitudinally extending passage in the tube 30 is in communication with the tap 13 for conveying hypergolic fluid into an enlarged pocket 31 formed in the end of the charge of solid propellant 19. A mass of solid propellant such as a disk 32 of plasticized polyvinyl chloride and ammonium perchlorate is positioned adjacent the tube 30 at the bottom of the pocket 31. The fluid is concentrated in the pocket 31 to react with both the solid propellant of the tube 19 and the disk 32 before it escapes through the nozzle 22 after the disk 32 burns.
The pocket type injector 23 was mounted in the apparatus shown in FIG. 1, and two tests were performed using cubic centimeters of chlorine trifluoride pressurized with 15 p.s.i.g. helium gas. The tube 30 had an inside diameter of 0.104 inch and extended into a one inch diameter pocket 31 for a distance of /a-inch. The disk 32 was positioned 1% inches from the head end of the charge 19, and the two tests produced reaction times of 0.31 and 0.34 second.
Referring to the embodiment of the invention shown in FIG. 3, the injector means 23a comprises a longitudinally extending tube 35 having its innermost end 36 closed.
A plurality of radial apertures 37 extend through the walls of the tube 35 to concentrate the hypergolic fluid. on the walls of the solid propellant charge 19.
A radial injector of the type shown in FIG. 3 comprising a tube 35 having an inside diameter of 0.104 inch I was tested in the apparatus shown in FIG. 1. This tube 35 extended into the bore for a distance of A-inch and} four radially-drilled holes 37, each having a diameter of 0.06 inch were provided for directing the fluid against the solid propellant. Hypergolic fluid in the storage bot- I tle 15 was pressurized to 15 p.s.i.g. and reaction times of 0.55 and 0.59 second were observed.
Referring now to the Y type injector means 23b shown in FIG. 4, hypergolic fluid is conveyed to the head end of the tube 19 of solid propellant through the tap 13 as apparatus shown in FIG. 1. This injector included a pair 1 of drilled passages each having a diameter of 0.072 inch that were located at an angle of 90 degrees from one another. A spacing of 0.313 inch was provided between the ends of the passages 40 and 41, and a reaction time of 0.71 second was observed when the hypergolic fluid in the bottle 15 was pressurized to 15 p.s.i.g.
The axial type injector 230 shown in FIG. 5 consists of an elongated tube 44 coaxial with the bore in the tube 19 of solid propellant. Such a tube 44 having an inside diameter of 0.104 inch was mounted on the apparatus shown in FIG. 1 to extend into the bore in the solid propellant charge 19, a distance of /2-inch. A reaction time of 0.91 second was observed when the hypergolic fluid was pressurized to 15 p.s.i.g.
While the data in the above examples Was obtained, using a measured quantity of 5 cubic centimeters of ignition fluid, a composite propellant motor was also ignited with 2.5 and 10 cubic centimeters of chlorine trifluoride. The ignition characteristics observed were approximately the same for all these fluid quantities.
Several illustrative embodiments of the invention have been disclosed and various structural modifications may be made without departing from the spirit of the invention or the scope of the subjoined claims. For example, the invention is not limited to any particular type of solid propellant, and the various types of injectors that have been disclosed as being used with composite solid propellants are equally applicable to double-base propellant motors. Also, the hypergolic fluids are not limited to the three particular aforementioned fluorides.
What is claimed is:
1. Apparatus for igniting a tube of solid propellant having a longitudinal bore extending therethrough, said apparatus comprising means at one end of said tube for directing a stream of hypergolic fluid along the axis of said bore, and
a member positioned in the path of 'said stream of hypergolic fluid for deflecting the flow thereof toward the wall of said bore and initially concentrating said fluid at said one end to react with said solid pro pellant in said tube before escaping from the opposite end of said bore.
. 2. Apparatus for igniting a tube of solid propellant having a longitudinal bore extending therethrough, said apparatus comprising injector means mounted at one end of said tube for introducing hypergolic fluid into said bore, said injector means comprising an elongated member extending into said bore at said one end thereof, said member having a longitudinally extending passage therein for conveying said fluid therethrough into said bore, and
a a mass of solid propellant positioned in said bore adjacent said elongated member for forming an obstruction which initially blocks the flow of said hypergolic fluid thereby concentrating said fluid at said one end to react with said solid propellant in said tube and said mass before escaping from the opposite end of said bore.
3. Apparatus for igniting a tube of solid propellant as claimed in claim 2, wherein the mass of solid propellant in the bore comprises a disc of plasticized polyvinyl chloride and ammonium perchlorate.
4. Apparatus for igniting a tube of solid propellant as in claim 1, wherein said member deflects the flow of the stream of hypergolic fluid at an angle relative to the longitudinal bore of said tube.
References Cited by the Examiner UNITED STATES PATENTS 2,685,168 8/54 Malick 6039.74 2,698,514 1/55 Hull et al. 6035.6 2,708,341 5/55 Zucrow 6039.06 2,850,975 9/58 Teague 6035.6 2,859,588 11/58 Wilson 6039.74 2,955,420 10/60 Schirmer 6039.06 2,972,227 2/61 Allen 60-39.45 2,974,484 3/61 Cooley 60--39.82 2,984,973 5/61 Stegelman 6035.6 2,990,682 7/61 Mullaney 6035.6 3,068,641 12/62 Fox 6035.4 3,128,599 4/64 Carr 6035.6 3,144,751 8/64 Blackman et al. 6035.6
SAMUEL LEVINE, Primary Examiner.
CARLTON R. CROYLE, Examiner.

Claims (1)

1. APPARATUS FOR IGNITING A TUBE OF SOLID PROPELLANT HAVING A LONGITUDINAL BORE EXTENDING THERETHROUGH, SAID APPARATUS COMPRISING MEANS AT ONE END OF SAID TUBE FOR DIRECTING A STREAM OF HYPERGOLIC FLUID ALONG THE AXIS OF SAID BORE, AND A MEMBER POSITIONED IN THE PATH OF SAID STREAM OF HYPERGOLIC FLUID FOR DEFLECTING THE FLOW THEREOF TOWARD THE WALL OF SAID BORE AND INITIALLY CONCENTRATING SAID FLUID AT SAID ONE END TO REACT WITH SAID SOLID PROPELLANT IN SAID TUBE BEFORE ESCAPING FROM THE OPPOSITE END OF SAID BORE.
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Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3274775A (en) * 1962-07-18 1966-09-27 Onera (Off Nat Aerospatiale) Hot gas generators and in particular rocket motors
US3286472A (en) * 1964-02-24 1966-11-22 Thiokol Chemical Corp Rocket motor and ignition system
US3295323A (en) * 1965-02-24 1967-01-03 Allen L Holzman Means for vaporizing liquid propellants
US3302403A (en) * 1964-06-25 1967-02-07 Leroy J Krzycki Method and apparatus for injecting a liquid propellant in hybrid rocket motors
US3315472A (en) * 1961-08-30 1967-04-25 Onera (Off Nat Aerospatiale) Hypergolic gas generator
US3334489A (en) * 1964-06-04 1967-08-08 Propulsion Par Reaction Sa Soc Rocket motor

Citations (13)

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Publication number Priority date Publication date Assignee Title
US2685168A (en) * 1948-01-02 1954-08-03 Phillips Petroleum Co Combustion chamber
US2698514A (en) * 1952-01-14 1955-01-04 Edwin H Hull Rocket motor
US2708341A (en) * 1951-07-25 1955-05-17 Kellogg M W Co Method of using a hot vapor igniter
US2850975A (en) * 1954-05-27 1958-09-09 Bendix Aviat Corp Acceleration pressurized bi-propellant liquid fuel rocket
US2859588A (en) * 1953-12-16 1958-11-11 Texas Co Flame holder structure for ram jet combustor
US2955420A (en) * 1955-09-12 1960-10-11 Phillips Petroleum Co Jet engine operation
US2972227A (en) * 1956-07-05 1961-02-21 Bristol Siddeley Engines Ltd Means for supplying a rocket motor with liquid fuel and concentrated hydrogen peroxide as propellant
US2974484A (en) * 1952-01-23 1961-03-14 Robert A Cooley Ignition system for rocket motors
US2984973A (en) * 1958-12-08 1961-05-23 Phillips Petroleum Co Liquid-solid bipropellant rocket
US2990682A (en) * 1951-11-07 1961-07-04 Gen Electric Fuel charge
US3068641A (en) * 1955-04-18 1962-12-18 Homer M Fox Hybrid method of rocket propulsion
US3128599A (en) * 1958-10-06 1964-04-14 Phillips Petroleum Co Solid-fluid rocket propellant system
US3144751A (en) * 1961-05-10 1964-08-18 United Aircraft Corp Hybrid rocket

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2685168A (en) * 1948-01-02 1954-08-03 Phillips Petroleum Co Combustion chamber
US2708341A (en) * 1951-07-25 1955-05-17 Kellogg M W Co Method of using a hot vapor igniter
US2990682A (en) * 1951-11-07 1961-07-04 Gen Electric Fuel charge
US2698514A (en) * 1952-01-14 1955-01-04 Edwin H Hull Rocket motor
US2974484A (en) * 1952-01-23 1961-03-14 Robert A Cooley Ignition system for rocket motors
US2859588A (en) * 1953-12-16 1958-11-11 Texas Co Flame holder structure for ram jet combustor
US2850975A (en) * 1954-05-27 1958-09-09 Bendix Aviat Corp Acceleration pressurized bi-propellant liquid fuel rocket
US3068641A (en) * 1955-04-18 1962-12-18 Homer M Fox Hybrid method of rocket propulsion
US2955420A (en) * 1955-09-12 1960-10-11 Phillips Petroleum Co Jet engine operation
US2972227A (en) * 1956-07-05 1961-02-21 Bristol Siddeley Engines Ltd Means for supplying a rocket motor with liquid fuel and concentrated hydrogen peroxide as propellant
US3128599A (en) * 1958-10-06 1964-04-14 Phillips Petroleum Co Solid-fluid rocket propellant system
US2984973A (en) * 1958-12-08 1961-05-23 Phillips Petroleum Co Liquid-solid bipropellant rocket
US3144751A (en) * 1961-05-10 1964-08-18 United Aircraft Corp Hybrid rocket

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3315472A (en) * 1961-08-30 1967-04-25 Onera (Off Nat Aerospatiale) Hypergolic gas generator
US3274775A (en) * 1962-07-18 1966-09-27 Onera (Off Nat Aerospatiale) Hot gas generators and in particular rocket motors
US3286472A (en) * 1964-02-24 1966-11-22 Thiokol Chemical Corp Rocket motor and ignition system
US3334489A (en) * 1964-06-04 1967-08-08 Propulsion Par Reaction Sa Soc Rocket motor
US3302403A (en) * 1964-06-25 1967-02-07 Leroy J Krzycki Method and apparatus for injecting a liquid propellant in hybrid rocket motors
US3295323A (en) * 1965-02-24 1967-01-03 Allen L Holzman Means for vaporizing liquid propellants

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